Analyzing Cessna 172 using DATCOM

Digital DATCOM^[1] is a tool to obtain aerodynamic characteristics of fixed wing aircrafts useful during preliminary design. However, its difficulty in use makes it prohibitive to beginners. To help getting started, this page provides an example DATCOM input file for Cessna 172 along with explanations of the different parameters. Python scripts to visualize both input and output files are provided. To check the correctness of DATCOM, its results were compared against those obtained via CFD^[2]. Below are lift, drag, and pitching moment coefficients obtained from DATCOM and CFD:

(Above was plotted with plot_result.py)

DATCOM Input File

The following DATCOM input file (cessna172.inp) was used:

CASEID Wing + Body + Aileron
DIM M
DERIV RAD
 $FLTCON
  LOOP=3.0,
  NMACH=1.0,
  MACH(1)=0.16,
  NALT=1.0,
  ALT(1)=1500.0,
  NALPHA=12.0,
  ALSCHD(1)=-7.5,-5.0,-2.5,0.0,2.5,5.,7.5,10.,15.,17.,18.,19.5$
 $SYNTHS
  XCG=2.11,
  ZCG=1.26,
  XW=1.68,
  ZW=2.04,
  ALIW=1.5,
  XH=6.16,
  ZH=1.68,
  ALIH=0.0,
  XV=5.88,
  ZV=1.55,
  SCALE=1.0,
  VERTUP=.TRUE.$
 $BODY
  NX=8.0,
  ITYPE=2.0,
  METHOD=2.0,
  X(1)=0.10,0.30,0.77,1.22,1.68,3.40,3.91,7.41,
  R(1)=0.14,0.32,0.49,0.52,0.72,0.59,0.44,0.10,
  ZU(1)=1.42,1.49,1.59,1.65,2.04,1.98,1.77,1.75,
  ZL(1)=1.14,0.85,0.61,0.61,0.61,0.80,0.90,1.55$
NACA-W-4-2412
 $WGPLNF
  CHRDR=1.626,
  CHRDBP=1.626,
  CHRDTP=1.13,
  SSPN=5.49,
  SSPNE=4.75,
  SSPNOP=2.941,
  SAVSI=0.0,
  SAVSO=3.0,
  CHSTAT=0.0,
  TWISTA=-3.0,
  DHDADI=1.7333,
  DHDADO=1.7333,
  TYPE=1.0$
 $ASYFLP 
  NDELTA=9.0,
  STYPE=4.0, 
  DELTAL(1)=-15.,-10.,-5.,-2.5,0.,5.,10.,15.,20.,                              
  DELTAR(1)=20.,15.,10.,5.,0.,-2.5,-5.,-10.,-15.,                          
  CHRDFI=0.312,
  CHRDFO=0.312,                                 
  SPANFI=2.44,
  SPANFO=5.15$
SAVE
NEXT CASE
CASEID Wing + Body + Vertical Tail + Horizontal Tail + Elevator
NACA-V-4-0009
 $VTPLNF
  CHRDR=1.53,
  CHRDTP=0.70,
  SSPN=1.62,
  SSPNE=1.42,
  SAVSI=33.0,
  CHSTAT=0.25,
  TYPE=1.0$
NACA-H-4-0012
 $HTPLNF
  CHRDR=1.24,
  CHRDTP=0.77,
  SSPN=1.73,
  SSPNE=1.50,
  SAVSI=9.0,
  CHSTAT=0.0,
  TWISTA=0.0,
  DHDADI=0.0,
  TYPE=1.0$
 $SYMFLP
  FTYPE=1.0,
  SPANFI=0.23,
  SPANFO=1.727,
  CHRDFI=0.56,
  CHRDFO=0.27,
  NTYPE=1.0,
  NDELTA=9.0,
  DELTA(1)=-26.0,-20.0,-10.0,-5.0,0.0,7.5,15.0,22.5,28.0$
DAMP

Below is a 3-view visualization of the above input file:

Side view (plotted with datcom_side_view.py):

Top view (plotted with datcom_top_view.py):

Front view (plotted with datcom_front_view.py):

CASEID

CASEID is a user defined title for a "case." DATCOM outputs aerodynamic characteristics for each case together in a single output file. cessna172.inp has two cases:

• Wing + Body + Aileron
• Wing + Body + Vertical Tail + Horizontal Tail + Elevator

The reason it was not done in a single case is because DATCOM did not output rolling moment coefficient due to aileron deflection when vertical and horizontal tails were present.

Units

• DIM M: units are in meters
• DERIV RAD: output static and dynamic derivatives in radians

FLTCON: Flight Conditions

• LOOP=3.0: 3 means vary altitude at fixed Mach number. We specify only a single altitude so only one study is done.
• NMACH=1.0: number of flight speeds that we run the analysis. We analyze at only one speed.
• MACH(1)=0.16: the flight speeds at which we analyze. We provide only one speed of 0.16 Mach which is about 198km/h or 55m/s.
• NALT=1.0: number of altitudes that we run the analysis. We analyze at only one altitude.
• ALT(1)=1500.0: the altitudes at which we analyze. We provide only one altitude of 1.5km.
• NALPHA=12.0: number of angles of attack that we run the analysis. We provide 12 angles in total.
• ALSCHD(1)=...: the angles of attack at which we analyze. We provide 12 angles from -7.5 to 19.5 degrees.

Aircraft Geometry

Aircraft geometry is defined in the following "namelists" (groups of parameters).

• SYNTHS: positions of center of gravity and wings
• BODY: body geometry
• WGPLNF: main wing geometry
• ASYFLP: aileron geometry
• VTPLNF: vertical tail geometry
• HTPLNF: horizontal tail geometry
• SYMFLP: elevator geometry

Wings and tails can be in two different shapes: straight or cranked. Straight means the wing is trapezoid. Cranked means there is an additional bend in the middle.

Coordinate system (according to DATCOM user manual SYNTHS section):

• +X: towards the back
• +Y: towards the right
• +Z: towards the top

Main Wing (WGPLNF) and Horizontal Tail (HTPLNF)

Top view of straight tapered planform:

Front view of straight tapered planform:

Top view of cranked planform:

Front view of cranked planform:

Vertical Tail (VTPLNF)

Side view of straight tapered planform:

Side view of cranked planform:

Aileron (ASYFLP) and Elevator (SYMFLP)

Top view of aileron:

Airfoil Shape

The following airfoil shapes were set:

• NACA-W-4-2412: main wing is NACA 2412
• NACA-V-4-0009: vertical tail is NACA 0009
• NACA-H-4-0012: horizontal tail is NACA 0012

DATCOM Output File

DATCOM output file (datom.out) is a text file. It is not shown here but can be downloaded from the files section. This file can be summarized as having the following 5 sections:

• Case: Wing + Body + Aileron
  • CHARACTERISTICS AT ANGLE OF ATTACK AND IN SIDESLIP --- A
  • CHARACTERISTICS OF HIGH LIFT AND CONTROL DEVICES --- B
• Case: Wing + Body + Vertical Tail + Horizontal Tail + Elevator
  • CHARACTERISTICS AT ANGLE OF ATTACK AND IN SIDESLIP --- C
  • DYNAMIC DERIVATIVES --- D
  • CHARACTERISTICS OF HIGH LIFT AND CONTROL DEVICES --- E

We are concerned with sections B, C, D, and E. Coefficients are summarized as follows:

Longitudinal Aerodynamics:

CD_total = CD + D(CDI)
CL_total = CL + D(CL) + c/(2*V) * (CLQ * q + CLAD * alpha_dot)
Cm_total = CM + D(CM) + c/(2*V) * (CMQ * q + CMAD * alpha_dot)

Lateral Aerodynamics:

CY_total = CYB * beta + b/(2*V) * (CYP * p)
Cl_total = CLB * beta + (CL)ROLL + b/(2*V) * (CLP * p + CLR * r)
Cn_total = CNB * beta + CN + b/(2*V) * (CNP * p + CNR * r)

Units for beta is in rad and p,q,r are in rad/sec as specified in DERIV control card.

Drag Coefficient (CD):

Drag increases with increasing angle of attack. Deflecting the elevator in either direction results in increased drag.

Lift Coefficient (CL):

Lift increases with increasing angle of attack. Deflecting the elevator down causes greater lift. Deflecting the elevator up causes downward force.

Pitching Moment Coefficient (Cm):

Increasing the angle of attack causes a negative pitching moment and decreasing angle of attack causes a positive pitching moment. This shows that it is stable. This is most likely because the center of gravity is in front of the aerodynamic center of the main wing. Deflecting the elevator down causes a negative pitching moment which means nose down. CMQ is -14.73, a negative number which means that it damps any pitching movement (aerodynamic damping).

Side-Force Coefficient (CY):

CYB is -0.2909, a negative number indicating that a positive sideslip angle causes a negative sideforce. This means that flying nose left causes a force to the left.

Rolling Moment Coefficient (Cl):

CLB is negative indicating a positive sideslip angle causes a negative rolling moment. This means that flying nose left causes the plane to roll to the left. Positive aileron deflection (right aileron down) causes negative rolling moment (roll left). CLP is negative indicating that rolling motion is damped out.

Yawng Moment Coefficient (Cn):

CNB is 0.02321, a positive number indicating that a positive sideslip causes a positive yawing moment. This means that flying nose left causes the plane to yaw to the right (it stabilizes). This is probably due to the vertical stabilizer. CN is mostly positive indicating that a positive aileron deflection (right aileron down) causes a positive yawing moment (yaw to the right). This opposes the motion of rolling to the left from the aileron and this probably arises as adverse yaw effect. CNP is mostly negative indicating that yaw motion is damped out.

Coefficients under "CHARACTERISTICS AT ANGLE OF ATTACK AND IN SIDESLIP"

• CD
  • Drag coefficient
  • Function of alpha
  • Positive for a force acting towards the back (positive for an aft-acting load)
  • Stability axis
  • Note that there is an equivalent in body axis called "CA" which is called axial-force coefficient which is defiend positive towards the back
• CL
  • Lift coefficient
  • Function of alpha
  • Positive for a force acting upwards
  • Stability axis
  • Note that there is an equivalent in body axis called "CN" which is called normal force coefficient defined positive up
  • Note that there is also "CLA" which is derivatives of lift coefficients relative to alpha
• CM
  • Pitching-moment coefficient
  • Function of alpha
  • Positive for a nose-up rotation
  • Stability axis
  • Note that there is also "CMA" which is derivatives of pitching-moment coefficients relative to alpha.
• CYB
  • Derivatives of side-force coefficients relative to sideslip angle
  • Function of alpha
• CNB
  • Derivatives of yawing-moment coefficients relative to sideslip angle
  • Function of alpha
• CLB
  • Derivatives of rolling-moment coefficients relative to sideslip angle
  • Function of alpha

Coefficients under "DYNAMIC DERIVATIVES"

• Pitching
  • CLQ
  • Lift force derivatives due to pitch rate
  • Function of alpha
  • CMQ
  • Pitching-moment derivatives due to pitch rate
  • Function of alpha
• Acceleration
  • CLAD
  • Lift-force derivatives due to rate of angle of attack
  • Function of alpha
  • CMAD
  • Pitching-moment derivatives due to rate of angle of attack
  • Function of alpha
• Rolling
  • CLP
  • Rolling-moment derivatives due to roll rate
  • Function of alpha
  • CYP
  • Lateral-force derivatives due to roll rate
  • Function of alpha
  • CNP
  • Yawing-moment derivatives due to roll rate
  • Function of alpha
• Yawing
  • CNR
  • Yawing-moment derivatives due to yaw rate
  • Function of alpha
  • CLR
  • Rolling-moment derivatives due to yaw rate
  • Function of alpha

Coefficients under "CHARACTERISTICS OF HIGH LIFT AND CONTROL DEVICES"

• CN
  • Function of the DELTAL-DELTAR (difference between left and right aileron deflection) and alpha
  • Under "YAWING MOMENT COEFFICIENT,CN,DUE TO CONTROL DEFLECTION"
• (CL)ROLL
  • Incremental rolling moment coefficient due to asymmetrical deflection of control surface
  • Positive rolling moment is right wing down
  • Function of aileron deflection angle
• D(CL)
  • Incremental lift coefficient in the linear-lift angle-of-attack range due to a deflection of control surface
  • Function of elevator deflection
• D(CM)
  • Incremental pitching-moment coefficient due to control surface deflection valid in the linear lift angle-of-attack range
  • Function of elevator deflection angle
• D(CDI)
  • Incremental induced-drag coefficient due to flap deflection
  • Function of elevator deflection and alpha
  • Under "INDUCED DRAG COEFFICIENT INCREMENT , D(CDI) , DUE TO DEFLECTION"

Nomenclature and Conventions

• alpha
  • Angle of attack
  • Positive nose up
• beta
  • Sideslip angle
  • Positive nose left
• Aileron
  • Positive right aileron down (usually causes the airplane to roll left)
• Elevator
  • Positive down deflection (usually causes the airplane to nose down)
• c
  • Mean aerodynamic chord length
• b
  • Wingspan (tip to tip)
• V
  • Airspeed
• Roll-pitch-yaw
  • Rotation along front-right-down
  • For example, positive rolling moment is right wing down

Files

DATCOM input and output files:

• cessna172.inp
• datcom.out

Files to visualize cessna172.inp:

• datcom_side_view.py
• datcom_top_view.py
• datcom_front_view.py
• datcom_inp_parser.py
• datcom_helper.py
• cessna172_side_view.png
• cessna172_top_view.png
• cessna172_front_view.png

Files to visualize datcom.out:

• plot_result.py
• datcom_out_parser.py
• cfd.csv

Note: above parsers are implemented to work for this specific case and not for any general DATCOM file so they may need modifications to work for different DATCOM cases.

Environment

Python scripts were run on Windows:

• Python 3.9.6
• numpy 2.0.2
• matplotlib 3.9.4
• pandas 2.2.3

DATCOM was compiled and run on Ubuntu:

• APRIL 1976 VERSION OF DATCOM (software was downloaded from here and the version number is noted in datcom.out file)
• GNU Fortran (Ubuntu 9.3.0-10ubuntu2) 9.3.0
• Ubuntu 20.04 LTS (on WSL2)

sudo apt-get install gfortran              # Install GNU Fortran compiler
gfortran -std=legacy datcom.f -o datcom    # Compile datcom.f (this generated many warnings)
./datcom                                   # Run DATCOM program

The last command will ask for an input file. After specifying cessna172.inp, it will generate datcom.out.

References

[1] AFFDL-TR-79-3032: The USAF Stability and Control DATCOM, Volume 1, User's Manual

[2] Cel, Marek. (2019). Cessna 172 Flight Simulation Data. 10.13140/RG.2.2.27040.51205.